Conformal head-up display

ABSTRACT

A system for providing a Head-Up Display (HUD) on board an aircraft to assist a pilot in guiding the aircraft. The display is positioned in the pilot&#39;s normal line of sight. In one mode it utilizes a radio beam landing system such as an ILS (Instrument Landing System) to generate symbols that correspond to visual ground cues which, together with an aircraft symbol display, provide the pilot with cues for aligning the aircraft on the appropriate path for approach and landing. The system moves the aircraft symbol in accordance with motion changes of the aircraft. During a landing approach the pilot &#34;flies&#34; the aircraft symbol relative to the simulated and/or real ground cues. By making the HUD correspond with ground cues, the abrupt transition from instrumented to visual flight is eliminated. At altitudes below which available ILS is not acceptable, the system provides a smooth transition to a mode independent of ILS and then to a flare mode. The system also includes a mode that is totally independent of any ground installation. The system also provides an altitude hold mode with a smooth transition to the approach mode.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to aircraft control systems and, moreparticularly, to systems which present to the pilot position-controlled,generated image, cue indications on a head-up display (HUD) superimposedwith visible external cues.

2. Description of the Prior Art

A head-up display (HUD) is one in which a generated image is projectedonto a transparent screen in the pilot's normal line of sight so thatthe pilot can observe the projected image while continuing his normalobservation of outside cues.

Prior HUD systems have been developed and are well known, such as alead-computing gunsight for military aircraft and approach-to-landingpath displays. The latter type of HUD designs hve made use of aselectable fixed approach path depression angle and/or the instantaneousdescent angle of the aircraft (velocity vector) either separately orcombined in one display symbol.

One HUD system, of particular interest for commercial aircraft, is thesystem disclosed in my prior U.S. Pat. No. 4,104,612 entitled HEAD-UPDISPLAY COMMAND BAR GENERATOR. In that disclosure, afeedback-compensated control system makes use of the fixed depressionangle combined with computed rate information to derive the HUD commandsignal. The pilot, by maneuvering to keep the symbol on an aimpoint (thedesired touchdown zone), closes the loop of a feedback-compensatedcontrol system. The system is referred to herein as the "CompensatedControl" HUD.

SUMMARY OF THE INVENTION

Arrangements in accordance with the present invention permit the pilotto fly the aircraft head-up during approach and landing for virtuallyall conditions which may be encountered by the aircraft. In accordancewith one aspect of the invention, the HUD is arranged so that the pilotis able to use the display in the same manner in both IFR (InstrumentFlight Rules) and VFR (Visual Flight Rules) conditions, thus avoiding ashift from one type of cue to another as the pilot proceeds from IFRconditions to VFR conditions. The system also provides a smoothtransition from an instrumented approach to a mode independent of theILS (Instrument Landing System) and then to a flare mode. Particulararrangements in accordance with the present invention providemagnification of the vertical control cue of the Compensated Control HUDin accordance with glideslope and glideslope rate information, derivedfrom an ILS beam, during approach for landing under such control.

The Compensated Control HUD does not require ground-basedinstrumentation to provide guidance to a runway that is in sight. Anarrangement in accordance with the present invention provides lateralguidance that is cued from a visible runway and requires no ground-basedinstrumentation. In this arrangement the aircraft symbol display iscontrolled to remain astride the runway which is being approached, eventhough the aircraft may be crabbed into a crosswind.

One particular arrangement in accordance with the present inventionserves to control the aircraft (at the pilot's option) at a selectedfixed altitude prior to pitching over to proceed on the selected descentpath. In this arrangement, the location of a reference aim symbolrelative to the aircraft symbol in the display provides an altitude holdcommand. The pilot by maneuvering his aircraft to hold the HUD aircraftsymbol on the reference symbol will cause the aircraft to maintain theselected altitude. When the aircraft symbol intercepts the runwayaimpoint (visible from the ground through the display), the pilot knowshe is at the point where he should begin his selected descent path andhe then pitches over to proceed toward the runway aimpoint.

BRIEF DESCRIPTION OF THE DRAWINGS

A better understanding of the present invention may be had from aconsideration of the following detailed description, taken inconjunction with the accompanying drawings in which:

FIG. 1 is an illustration of one particular HUD display in accordancewith the invention, showing runway parameters and other generated flightcues projected on the display;

FIG. 2 is a block diagram of circuitry for incorporating glideslopeerror signals to amplify the aircraft symbol vertical deflection signalin an ILS approach;

FIG. 3 is an illustration of the HUD parameters related to aninstrumented approach;

FIGS. 4A and 4B illustrate the geometric parameters required in thecomputation of the display parameters of FIG. 3;

FIG. 5 is a block diagram illustrating circuitry for determining lateraldeflection of the aircraft symbol as shown in FIG. 3 in accordance withthe geometric parameters of FIGS. 4A and 4B;

FIGS. 6A and 6B illustrate in the display and the vertical plane theparameters related to the preflare and flare modes;

FIG. 7 is a block diagram of circuitry for controlling the aircraftsymbol vertical deflection in the final phases of landing approach;

FIG. 8 is a block diagram of circuitry for controlling the aim symbolvertical deflection in the final phases of landing approach;

FIG. 9 is a block diagram illustrating particular filter circuitry toprovide lateral command with lead for a non-instrumented approach;

FIGS. 10A and 10B illustrate aircraft locations and correspondingdisplays during altitude hold and transition to the approach;

FIG. 11 is a block diagram of circuitry for controlling verticaldeflection of the aim symbol in the altitude hold mode;

FIG. 12 is a block diagram illustrating particular lead circuitry forcontrolling vertical deflection of the aircraft symbol to providealtitude hold damping and transference to the descent path; and

FIG. 13 is a schematic diagram illustrating a system for generating ahead-up display in accordance with the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

The Compensated Control HUD of my prior patent provided verticalguidance for the pilot during descent prior to landing. Preferredembodiments of the present invention utilize my patented CompensatedControl HUD and in addition provide HUD lateral and vertical controlguidance in a manner which minimizes the transition from IFR to VFRconditions. It also provides command guidance from the minimum ILScontrol altitude to flare, during a non-instrumented approach and duringaltitude hold including a smooth transition to a descent path. The ILSis referenced herein, but the invention is not restricted to anyparticular radio beam landing system.

My Compensated Control HUD is disclosed in detail in my prior U.S. Pat.No. 4,104,612, incorporated by reference herein, and the reader isreferred to that patent for a more detailed discussion of that systemand a showing of the image generator and projection arrangement whichdevelops the HUD. Summarizing very briefly from that patent, a selectedglidepath angle plus aircraft motion terms like pitch, altitude rate andnormal acceleration are coupled into the control law circuit module,which generates the deflection signal. This deflection signal is coupledinto a symbol generator which generates the image of a command bar(aircraft symbol) on a cathode ray tube (CRT). The deflection signalcontrols the vertical location of the command bar. The CRT image isprojected onto a collimating lens and to a combiner, which is anon-interfering transparent glass located between the pilot's eyes andhis forward field of view. The command bar appears to the pilot assuperimposed on his view of the runway during a landing approach. Thecollimated command bar constitutes an aimsight. In steady state, adeflection signal transmitted from the control law module represents theselected glidepath angle, causing the aimsight (collimated command bar)to be held at the selected angle relative to the horizontal. If theaircraft is on the selected glidepath, the pilot will see the barsuperimposed on his runway aimpoint. Should the pilot see the bar aboveor below his aimpoint, he interprets it as control error and correctswith elevator control. An aircraft maneuver produces changes in theaircraft motion terms, which are coupled into the control law module,which produces a change in bar deflection, thus completing the loop ofthe Compensated Control HUD system.

One principal embodiment of the present invention is an augmentation tomy prior U.S. Pat. No. 4,104,612 which utilizes the glideslope signal ofa radio beam landing system such as an instrument landing system (ILS)to provide path error magnification. Though my prior patented system iscapable of providing vertical guidance in a VFR, non-instrumentedapproach to the accuracy required of an autopilot by FAA Circular AC120-29 for an IFR Category II approach, the addition of the ILSaugmentation of this embodiment will command closer tracking. Accordingto my prior patent the pilot is guided by the vertical motion of asymbol, herein called aircraft symbol 10 in FIG. 1, relative to a pointon the ground, herein called aimpoint 18 in FIG. 1. The verticallocation of the aircraft symbol relative to the aimpoint is determinedby the compensated control command signal. Where an ILS glideslope beamis available, the aimpoint can be supplemented with a HUD symbol, hereincalled aim symbol 14 in FIG. 1. Also with the aid of the glideslopeerror signal the aircraft symbol vertical command signal is amplified.This amplification increases the separation between the aim symbol 14and the aircraft symbol 10 for a given separation of the aircraft fromthe glidepath. By increasing the visibility of the error the pilot mayfly a tighter control during this instrumented approach.

In this embodiment vertical position of the aircraft symbol in thedisplay is determined by the Compensated Control law of my prior patentreferenced above plus the glideslope signal augmentation. A verticalsignal of the prior patent (η_(c) -θ+η_(1d)) is combined with ILSglideslope error and glideslope error rate as shown in FIG. 2. FIG. 2shows a Compensated Control circuit 13 from my prior patent combinedwith glideslope error proportional gain 15 and glideslope error ratederived from the rate taker 17. The rate taker time constant τ_(G)should be fast relative to aircraft pitch dynamics. Gains K_(G) andK_(G) are adjusted for optimum control. The outputs from stages 15 and17 are added to the Compensated Control output in a summing stage 19.When no ILS signal is available, the constants K_(G) and K_(G) are setto zero.

Another aspect of the invention relates to a method of combining thelateral flight director, vertical guidance, and a perspective display ofground cues so as to provide an instrumented (radio beam) guidance thatis conformal with the real world visual cues. The pilot may "fly" theaircraft symbol, which pitches, rolls and yaws relative to the realworld, to either a simulated runway aimpoint (a point at the center ofthe runway where the glideslope intersects the runway) while flyingduring restricted visibility or the actual runway when it becomesvisible.

FIG. 1 shows such a HUD comprising an aircraft symbol 10, a perspectiveline 12 that superimposes on the actual runway centerline when it isvisible, and an aim symbol 14 that appears like a line on the groundthat crosses the runway 30, 32, 34 parallel to the threshold 34 at theaimpoint, which is located at the glideslope antenna. The runwayaimpoint 18 is indicated by the intersection of the perspectivecenterline 12 and the aim symbol 14. (The aimpoint can be simulated byother symbologies; for example a perspective display of the total runway30, 32, 34 with aimpoint, or a small circle located at the aimpoint.)The horizon line 24 has displayed on it a line 50 indicating theaircraft heading and a "V" symbol 40 denoting the runway bearing, avalue which is set into the equipment by the pilot in preparation forthe approach. Thus, the symbol 40 relative to symbol 50 represents therunway bearing relative to aircraft heading, or Course Error. Theaircraft symbol 10 symbolizes an aft view of an aircraft controlled bythe pilot in pitch, roll and yaw to correspond to motions of hisaircraft. These motions are made evident by a pair of lines or "wings"16 emanating from a central point, which can be indicated by the circleas shown. During the approach the pilot "flies" the aircraft symbol tokeep it centered on the aimpoint 18. The particular symbology describedherein is only one example for picturing the space location of aircraftand ground cues in the display. A portion of the embodiment disclosedencompasses the generation of electronics that determine this spacelocation in the conformal display.

FIG. 3 illustrates parameters required in the generation of theconformal display for an instrumented approach. The vertical deflectionof the aircraft symbol 10 relative to the horizon 24 is determined bythe control law of my prior patent (Compensated Control law) and equalsη_(c) +η_(1d) where η_(c) equals the desired angle of descent and η_(1d)is the feedback compensation lead term. The Compensated Control circuitis designed so that if the pilot maneuvers his aircraft in the verticalplane to hold the aircraft symbol on some line fixed relative to theground, the aircraft will fly a path that descends at η_(c) degrees andterminates at the fixed point. The vertical deflection of th aircraftsymbol relative to the runway aimpoint is η_(e). If the aircraft descentangle is controlled to hold the center of the aircraft symbol on theaimpoint 18, the command is satisfied and the aircraft will smoothlyacquire and track the desired descent path. Aim symbol 14 is computed tostay fixed in the direction of the glideslope antenna on the ground.Therefore, by holding the aircraft symbol 10 on line 14, the pilot willfollow an η_(c) descent path that terminates at the glideslope antenna.

The lateral deflection of the aircraft symbol 10 relative to the runwayaimpoint 18 is generated from a signal that is like a lateral flightdirector roll command signal. It contains raw localizer signals pluslateral lead inputs. If the pilot maneuvers his aircraft laterally tohold the center of the aircraft symbol 10 on the aimpoint 18, the flightdirector command is satisfied and the aircraft will smoothly acquire andtrack the localizer beam.

The aim symbol 14 is located in the display at the runway aimpont bydepressing the symbol below the horizontal by an angle equal to thedepression angle of the aimpoint located at the glideslope transmitterantenna η_(AIM) (FIG. 4A). This is computed from the glideslope angleη_(c) plus the vertical angle deviation of the aircraft from theglideslope η_(GS) as measured by the glideslope error signal at theaircraft.

    η.sub.AIM =η.sub.c +η.sub.GS

Referring to FIG. 3, λ_(ca), the lateral deflection of the aimpoint 18relative to the display center (aircraft heading) 50, is equal to λ_(c),the horizontal angle between aircraft heading and selected runwaybearing (Course Error), plus λ_(a), the horizontal angle between runwaybearing and aimpoint.

    λ.sub.ca =λ.sub.c +λ.sub.a

λ_(a) is the product of the depression angle of the aimpoint η_(AIM) andthe ratio of the pilot's eye height (HEYE, FIG. 4A) to the lateraldistance his eye position is offset from an extension of the runwaycenterline (OSANT, FIG. 4B).

    λ.sub.a =η.sub.AIM ×HEYE/OSANT

The perspective line 12 is drawn from the selected runway bearing 40 onthe horizon to the aimpoint 18.

The following geometric computations are necessary to the solution ofthis embodiment. Referring to FIG. 4B, the distance the aircraftlocalizer receiver antenna is offset from the runway centerline OS iscomputed from the angle derived from the localizer error signal τ_(LOC)and the distance from the localizer transmitter R_(LOC). HEYE and OSANTare corrected for eye location relative to radio altimeter and aircraftlocalizer receiver antenna respectively. Finally the distance to thelocalizer transmitter is computed as the distance to the glideslopetransmitter R_(AIM), which is a function of altitude and glideslopedepression angle (described above), plus an approximate distance betweenglideslope transmitter and localizer transmitter R_(GL).

    R.sub.LOC =R.sub.AIM +R.sub.GL

At R_(AIM) distances that are less than the minimum limit of glideslopereliability R_(AIMIN), R_(LOC) becomes a function of airspeed TAS andtime t.

    R.sub.AIM <R.sub.AIMIN

    R.sub.LOC =R.sub.AIMIN +R.sub.GL -(TAS×t)

Runway edgelines, which may be added to the display, are computed in thesame manner as the centerline by adding and subtracting to OSANT therunway half width. The depression angle of the threshold η_(THR) andrunway far end can be computed as the arctangent of (HEYE/horizontaldistance from aircraft to the respective runway ends). The horizontaldistances are computed by adding or subtracting the distance betweenglideslope transmitter and runway ends.

FIG. 5 is a schematic representation in block diagram form of circuitrythat generates the signal λ_(ca), which locates in the display theperspective direction of the aimpoint as seen by the pilot, as describedhereinabove by reference to FIGS. 3 and 4A-4B. As is shown in FIG. 5,signals indicating pitch angle and altitude are applied to a stage 100which corrects altimeter reading to eye height, which is a function ofpitch angle, and provides a signal HEYE to divider stages 102 and 104.In the divider stage 102, HEYE is divided by a signal derived from thesummation of η_(c) and η_(GS) (see FIG. 4A). The output of divider 102,R_(AIM), is added to the signal R_(GL) to develop the signal R_(LOC)which is applied to one pole 106 of a switch 108. The armature 110 ofthe switch 108 is connected to a multiplier stage 112, the other inputto which is the localizer error signal τ_(LOC). The output of themultiplier 112 is the signal OS which is applied to a correction stage114 for combination with a signal λ_(c) which is the difference betweenaircraft heading and selected runway bearing. The correction stage 114provides the correction for lateral offset between the aircraft'slocalizer receiver antenna and the pilot's eyes when the aircraft is noton heading, and provides an output signal OSANT which is applied to thedivider 104. The output of divider 104, OSANT/HEYE, is multiplied byη_(AIM) in a multiplier stage 116 and the product, λ_(a), is summed withthe horizontal angle λ_(c) to provide the signal λ_(ca).

When the distance to the glideslope transmitter, R_(AIM), becomes lessthan the minimum limit of glideslope reliability, R_(AIMIN), thearmature 110 of switch 108 transfers to the pole 107 and R_(LOC) isgenerated as a function of time. Pole 107 receives a signal which is thedifference between R_(AIMIN) +R_(GL) and the integration (provided inintegrator stage 118) of true air speed. Under these conditions, thesignal at pole 107 is applied through switch 108 via armature 110 and isprocessed as previously described in place of the previous R_(LOC)signal.

Where it is considered unsafe to continue to follow the ILS glideslopebelow a minimum height, the pilot must either discontinue the landing orbe dependent upon runway and terrain visual cues. One preferredembodiment allows a continuation of HUD guidance below this minimumheight by transitioning from an instrumented approach to a modeindependent of the ILS and then to a flare mode. This embodiment alsoprovides smooth transitioning from one mode into the next.

A landing approach can be continued below ILS usability, with groundcues not visible, by maintaining the same angle of descent η_(c). Thisis accomplished by converting the aim symbol 14 so that it will indicatethe direction that it is desired to descend and converting the aircraftsymbol 10 so that it will indicate the direction the aircraft isdescending at any moment. The aim symbol, fixed relative to the ground,will become the command aim symbol 64 (FIG. 6A), which is depressedbelow the horizontal by the fixed angle η_(c) ; and the aircraft symbolwill be depressed below the horizontal by the velocity vector γ. Thus bymaneuvering to hold the aircraft symbol 10 on the aim symbol 64, theaircraft will be flown in the directon η_(c). In this mode the aimsymbol at η_(a) equals η_(c) and the aircraft symbol at η_(sym) equalsγ, the velocity vector.

The transference from the ILS dependent mode to the constant directionmode should be done at the lowest altitude that the ILS glideslope issatisfactory (designated "transition height"), because the constantdirection is maintained open loop and is dependent upon the accuracy ofobtaining γ, the velocity vector. In the absence of an InertialNavigation System, γ is dependent upon air measurements which areaffected by wind.

FIGS. 6A and 6B illustrate particular parameters used in preflare andflare modes. In the flare mode aim symbol 64 is converted to command aprogrammed sink rate reduction, while the aircraft symbol 10 isconverted to indicate a sink rate error relative to the aim symbol. Thusin this mode the aim symbol 64 at η_(a) becomes sink rate command andthe aircraft symbol at η_(sym) indicates sink rate. The flare embodimentof this invention, though described as associated with an instrumentedIFR approach, operates the same in a non-instrumented approach under VFRconditions.

An instantaneous change in mode at a selected altitude would produce astep in the input and a jump in the symbology. Circuitry for avoidingthis is depicted in the block diagrams of FIGS. 7 and 8. The circuit ofFIG. 7 controls the vertical deflection of the aircraft symbol 10 whilethe circuit of FIG. 8 controls the vertical deflection of the aim symbol64.

FIG. 7 depicts the circuitry that transfers aircraft symbol control fromCompensated Control to velocity vector by linearly reducing the gain ofthe Compensated Control input from one at transition height TH to zeroat flare height FH, while the gain of the velocity vector input islinearly increased from zero to one. This is depicted respectively at150 and 152. Then between flare height and touchdown, aircraft symbolcontrol is similarly transferred from velocity vector to altitude ratefeedback as shown by 154 and 156. The switching is indicatedconceptually by the switch 160. Thus, the switch 160 may be consideredto be at the position A for altitudes above TH (transition height), atthe position D between transition height and flare, and at the positionF for altitudes below the flare height.

FIG. 8 depicts the aim symbol position control that corresponds to theaircraft symbol modes described. Above TH the aim symbol designates thedirection of the aimpoint (or the glideslope antenna). At TH the gain ofthe ILS glideslope error signal input to the aim symbol begins to reducelinearly, reaching zero at flare height as indicated by 162, so that atflare height the aim symbol depression angle equals η_(c). Then betweenflare height and touchdown the aim symbol depression angle is reduced toan angle consistent with the desired touchdown descent rate. This isdepicted at 164 and 166. Mode switch 170 in FIG. 8 corresponds to 160 inFIG. 7.

Note that the switching altitudes described need not be restrictive. Forinstance in mode D, designated as between TH and flare, the gain changecould have been completed above flare and constant gain maintained untilflare height is reached. Also the gain changes need not be linear.

In FIG. 7 the velocity vector input to 152 and 154 is shown computed asvertical speed h divided by horizontal speed V. Horizontal speed may beobtained from airspeed (TAS) or ground speed where available. Leadinputs θ and A_(n) are combined with the velocity vector to make itflyable.

Flare is controlled according to aircraft altitude and descent rate. Theaim symbol in FIG. 8 converts from a descent angle reference η_(c) to adescent rate reference TD h CMD, while the aircraft symbol in FIG. 7converts from a measured descent angle to a measured descent rate. Thusthe vertical position of the aircraft symbol relative to the aim symbolcues the pilot that he is above or below a programmed reduction indescent rate as he approaches touchdown.

A further embodiment of the invention provides lateral guidance withoutthe aid of ground-based instrumentation such as the ILS. This, combinedwith my prior U.S. Pat. No. 4,104,612, which has been referred to hereinas the Compensated Control HUD, makes it possible to "fly" an aircraftsymbol both laterally and vertically to an aimpoint on a visible runwaywhere no ILS facilities are available. The lateral guidance aids thepilot in acquiring and holding a path that is colinear (lines up) withthe runway centerline. Lead information is incorporated in the lateraldeflection of the aircraft symbol to aid the pilot in anticipatingovershoot. Additionally, because the aircraft symbol is held astride therunway, vertical control to the runway aimpoint is made easier.

FIG. 9 shows the generation of the signal controlling the lateraldeflection λ_(VFR) of the aircraft symbol relative to aircraft heading(center of the display). In the short term the aircraft symbol moveslaterally with aircraft heading. In the steady state, however, thelateral position is washed out to selected runway heading. In steadystate the aircraft symbol lateral position is the angular distance ofrunway bearing (selected heading) from aircraft heading; this equals theCourse Error input. A change in aircraft heading is transmitted throughthe wash-out filter 182, 184 to cause, in the short term, the aircraftsymbol to hold its position relative to the display center, since equaland opposite signals are summed at 180. At the same time the real worldrunway moves in the direction opposite to the aircraft's turn (thusproviding lead relative to the real world). In the long term the signalthat is transmitted through the wash-out filter circuitry reduced tozero, so that again the aircraft symbol is deflected from the center ofthe display by the Course Error, and thus it remains in the direction ofthe runway bearing. The lateral deflection of the aircraft symbolrelative to aircraft heading is ##EQU1## The time constant ₁₀₅ τ of thewash-out circuit should be long relative to aircraft lateral dynamics(10 seconds for large aircraft). This design, as indicated by FIG. 9,produces lateral lead in a turn by the motion of the aircraft symbolrelative to the ground. The lead input to the lateral guidance may beobtained by various techniques. Since the purpose of the wash-outcircuit is to provide a deflection of the aircraft symbol in the shortterm for changes in aircraft heading, aircraft heading may besubstituted for the Course Error input to the wash-out circuit. Thisavoids the transient due to the change in Course Error when the pilotselects a new Selected Heading. ##EQU2## Additional lead may be gainedwith a roll angle input, with K.sub.Ψ 182 and K.sub.φ 186 gainsadjusted. Also if an Inertial Navigation System is available, theinertial lateral motion signal can provide the best possible controllead.

A further major embodiment comprises an altitude hold mode. In thealtitude hold mode, the aim symbol 64 of FIG. 6A provides a commandrelative to the aircraft symbol 10 that enables the pilot to hold hisaircraft at a constant selected altitude. It also enables him to fly asmooth transfer without overshoot to the descent path when the properpoint is reached without change in symbology or manner of control. A HUDapproach generally involves flying at a constant altitude while watchingthe HUD aircraft symbol as it moves relative to the ground. When thissymbol reaches his runway aimpoint, the pilot knows he is to pitch overand begin his descent along the selected descent path. This situation isrepresented in FIGS. 10A and 10B which show the altitude hold mode atpoints a through c. At d, the aircraft symbol 10, located below thehorizon by the fixed descent angle η_(c) is superimposed over theaimpoint. Thus at point d, the pilot pitches over and proceeds throughthe point e along the descent path to touchdown at the aimpoint. In thismanner altitude hold guidance and guidance for transitioning to landingapproach are provided as the pilot continues to fly head-up (lookingthrough the windshield).

The altitude hold mode comprises the two display symbols shown in FIG.10B, a guidance symbol, or aircraft symbol 10, and a command symbol, oraim symbol 64. While the aircraft holds the selected altitude theaircraft symbol and aim symbol will be located together at a fixeddepression angle η_(c) as shown at points c and d, FIGS. 10A-10B. Shouldthe aircraft be maintaining a constant altitude but below the selectedaltitude, the aim symbol will appear above the aircraft symbol by theangle η_(e) that is proportional to the altitude error. This is shown ata, FIGS. 10A-10B; this tells the pilot to "fly up". The pilot's task isto maneuver his aircraft to hold the aircraft symbol in line with theaim symbol as at b, c, and d. He "flies" the aircraft symbol to the aimsymbol. At b, as the aircraft climbs, the command is satisfied throughthe selected altitude has not yet been reached. Dynamic lead termsη_(ld) added to the fixed angle η_(c) have deflected the aircraft symbolup. This constitutes feedback damping, which aids the pilot in smoothlyacquiring and tracking the selected altitude.

The circuit of FIG. 11 serves to provide a command signal for aim symboldeflection. The selected altitude η_(SEL) is set into the equipment bythe pilot and is summed with the radio altimeter reading h in a summingstage 202. The resulting difference is applied through a gain stage 204which converts an altitude error to an aim symbol deflection angle; thegain is optimized for ease of tracking. When in the altitude hold mode,switch 206 is closed and the negative of the scaled altitude error isapplied to a summing stage 203 for combination with the descent pathangle η_(c) and the pitch angle θ. θ provides pitch stabilization (thatis, symbology will not pitch with aircraft), a requirement in any HUDsystem. The signal out of summing stage 203 is applied to the deflectioncircuit for the aim symbol 64.

The circuit of FIG. 12 serves to provide a command signal for aircraftsymbol deflection. The symbol is deflected by dynamic lead signalscomprising washed-out pitch, pitch rate, and washed-out altitude ratecoupled through optimized gains 209, 211 and 213 combined at summingstage 214. Other lead combinations can be used for altitude holddamping; for example washed-out pitch and pitch rate alone would beadequate. Also any lead η_(ld) disclosed in my prior Pat. No. 4,104,612would be satisfactory. With switch 206A in the position show (Alt.hold/Approach) integrator 212 will wash out the signal at summing stages207 according to time constant τ_(a). τ_(a) is made sufficiently long todamp long period altitude transients. Like the aim symbol deflection,the steady state aircraft symbol deflection is η_(c), which is pitchstabilized by φ at summing stage 208. The signal out of 208 is combinedwith the lead signals at summing stage 214 to produce the signal appliedto the deflection circuit for the aircraft symbol.

An important adjunct of the altitude hold function is the transferencefrom the altitude hold mode to the approach mode. FIGS. 10A-10B show theaim symbol 64 and the aircraft symbol 10 maintained in alignment at thefixed deflection angle η_(c) at point c as the aircraft approaches therunway 30, flying at a constant selected altitude. At point d thedisplay shows the aircraft symbol directly in line with the aimpoint onthe runway, thus signifying that the aircraft is on the selectedglidepath η_(c) where the pilot should be controlling the aircraft todescend. Switch 206 (FIG. 11) should be opened to remove the altitudehold signal from the aim symbol. The pilot may transfer his controlreference from the aim symbol to the aimpoint, thus transitioning to theapproach mode with negligible eye movement or attention change. Byremoving the altitude hold error signal from the vertical deflection ofthe aim symbol, the aim symbol is converted to the fixed depressionangle η_(c). The aim symbol may or may not be removed from the display.In IFR conditions, the aim symbol can be converted to designate therunway aimpoint 34.

Capture of the glidepath without overshoot may be made easier byreducing the time constants of the pitch and altitude rate wash-outsduring the period of capture as indicated in FIG. 12. At the instant ofpitch-over, 206A is switched to the Capture mode so that the integratorgains will be increased to K_(F) 210 and the time constant of thewash-outs will be reduced by the factor 1/K_(F). This allows the pitchand altitude rate steady state references to change more quickly fromthe values acquired in altitude hold to the glideslope values, thusminimizing the potential for overshoot. When the glidepath isestablished, 206A is switched back to Alt. Hold/Approach so that thetime constants will revert to the longer period to damp long periodaltitude rate transients during the approach. Further details as tovariations possible for aircraft symbol (command bar) lead circuitry ηldare described in my prior patent 4,104,612 and incorporated herein.

FIG. 13 depicts in block diagram form an overall system in accordancewith the present invention. As shown in FIG. 13, a cathode ray tube 220is used to develop a visual image which is applied through a collimatinglens 222 to a combiner 224 which presents the display seen by the pilot.The combiner is a non-interfering transparent glass located between thepilot's eyes and his forward field-of-view. The image of the cathode raytube 220 is developed by symbol generators 226 which generate thevarious symbols to be displayed in response to inputs from the controlcircuitry 230 as described hereinabove. Thus, the pilot is enabled toclose the loop between the system signal outputs generated by circuitry230 and his control of the aircraft by "flying" the aircraft symbol tocoincide with the aim symbol. Additional symbology provide referenceinformation to the pilot. Use of the system in this fashionsubstantially eases the pilot's task, particularly during landingapproaches under adverse weather conditions, and may permit landing ofthe aircraft under IFR conditions.

Although there have been described above specific arrangements of aconformal head-up display in accordance with the invention for thepurpose of illustrating the manner in which the invention may be used toadvantage, it will be appreciated that the invention is not limitedthereto. Accordingly, any and all modifications, variations orequivalent arrangements which may occur to those skilled in the artshould be considered to be within the scope of the invention as definedby the annexed claims.

What is claimed is:
 1. Apparatus for providing guidance information toan aircraft pilot in the form of a conformal head-up display (HUD) whichcomprises:means for generating an aircraft symbol on the display thatsimulates the attitude, location and motion of his aircraft relative tothe earth; means for generating symbols that simulate on the displayselected ground cues in position and orientation which superimpose overactual selected ground cues as seen by the pilot when they are visiblethrough the display; means for generating an aim symbol that ispositioned on the display to provide a reference to which the aircraftsymbol is directed for proper operation of the aircraft in selectedflight modes; means for controlling the position of the aircraft symboland the aim symbol in response to selected signal inputs to providevertical guidance information in different phases of a landing approachfrom above transition height, through transition to flare height, andthence to touchdown; means for switching control of the aircraft symboland of the aim symbol from one phase to the next; and means forproviding a smooth transition from one set of selected signal inputs tothe next as the aircraft descends through the respective phases.
 2. Theinvention according to claim 1 wherein the controlling meanscomprises:means for depressing the aim symbol by a fixed angle below thehorizontal, which depression angle may be selected by the pilot to equalthe glideslope radio beam angle at a runway being approached; and meansfor positioning the aircraft symbol to be at any instant depressed belowthe horizontal by an angle equal to the angle at which the aircraft isdescending at that instant, whereby by maneuvering the aircraft so thatthe aircraft symbol appears at the same depression angle as the aimsymbol the pilot will cause the aircraft to descend at the selecteddescent angle.
 3. The invention according to claim 1 including aircraftsymbol vertical deflection circuitry which comprises:means forgenerating a selected angle signal equivalent to the desired approachdescent angle; means for generating a feedback compensated controlsystem dynamic lead signal; summing amplifier means for combining saidselected angle signal with said dynamic lead signal to provide acompensated control system signal to control the vertical position ofthe aircraft symbol above transition height; means for generating belowtransition height a velocity vector signal which at any instant isequivalent to the angle of descent of the aircraft at that instant; andmeans for generating below flare height a descent rate signal that isequivalent to the measured descent rate of the aircraft.
 4. Theinvention according to claim 4 wherein the transition providing meanscomprises:first variable gain amplifier means to vary the amplitude ofsaid dynamic lead signal; means to reduce the gain of the first variablegain amplifier means from one above transition height to zero asaircraft altitude decreases; second variable gain amplifier means tovary the amplitude of said velocity vector signal; means to increase thegain of the second variable gain amplifier means from zero abovetransition height to one as the aircraft altitude decreases; and summingamplifier means for combining signals from the first and second variablegain amplifier means to provide a signal between transition height andflare height that smoothly transitions control of the aircraft symbolvertical deflection from dynamic lead signal control to velocity vectorsignal control.
 5. The invention according to claim 3 wherein thetransition providing means comprises:first variable gain amplifier meansto vary the amplitude of said velocity vector signal; means to reducethe gain of the first variable gain amplifier means from one at flareheight to zero as altitude decreases; second variable gain amplifiermeans to vary the amplitude of the descent rate signal; means toincrease the gain of the second variable gain amplifier means from zeroat flare height to one as altitude decreases; and summing amplifiermeans for combining the signals from the first and second variable gainamplifier means to provide, between flare height and touch-down, asignal that smoothly transitions control of the aircraft symbol verticaldeflection from velocity vector signal control to descent rate signalcontrol.
 6. The invention according to claim 1 including aim symbolvertical deflection circuitry which comprises:means for receiving aglideslope radio directional beam signal and developing a glideslopeerror signal indicative of vertical angular deviation of the aircraftrelative to the beam center; means for generating a selected descentangle signal corresponding to a selected approach angle of descent;summing amplifier means for combining the glideslope error signal withthe selected descent angle signal to provide a glideslope referencesignal to control the vertical position of the aim symbol abovetransition height; means for controlling the aim symbol verticaldeflection in response to the selected descent angle signal belowtransition height; and means for generating below flare height a descentrate command signal.
 7. The invention according to claim 6 wherein thetransition providing means comprises:variable gain amplifier means tovary the amplitude of the glideslope reference signal; means to reducethe gain of the variable gain amplifier means from one at transitionheight to zero as a function of decreasing aircraft altitude; andsumming amplifier means for combining the signal from the variable gainamplifier means with the selected angle signal to provide, betweentransition height and flare height, a signal that smoothly transitionscontrol of the aim symbol vertical deflection from glideslope referencesignal control to selected angle signal control.
 8. The inventionaccording to claim 6 wherein the transition providing meanscomprise:first variable gain amplifier means to vary the amplitude ofsaid selected descent angle signal; means to reduce the gain of thefirst variable gain amplifier means from one at flare height to zero asaircraft altitude decreases; second variable gain amplifier means tovary the amplitude of said descent rate command signal; means toincrease the gain of said second variable gain amplifier means from zeroat flare height to one as aircraft altitude decreases; and summingamplifier means to combine the signals from the two variable gainamplifier means to provide, between flare height and touch-down, asignal that smoothly transitions control of the aim symbol verticaldeflection from the selected descent angle signal control to the descentrate command signal control.
 9. The invention according to claim 1further including means for combining signal indicative of aircraftpitch rate and normal acceleration and applying a resultant signal tothe aircraft symbol positioning means in order to add dynamic lead tothe aircraft symbol motion to prevent aircraft symbol guidance fromcausing overshoot and oscillation in aircraft control.
 10. Apparatus forproviding guidance information to an aircraft pilot in the form of aconformal head-up display (HUD) which comprises:means for generating anaircraft symbol on the display that simulates the attitude, location andmotion of his aircraft relative to the earth; means for generatingsymbols that simulate on the display selected ground cues in positionand orientation which superimpose over actual selected ground cues asseen by the pilot when they are visible through the display; means forgenerating an aim symbol that is positioned on the display to provide areference to which the aircraft symbol is directed for proper operationof the aircraft in selected flight modes; and means for controlling theposition of the aircraft symbol and the aim symbol in response toselected signal inputs to provide vertical guidance information forconstant altitude control, said controlling means comprising: (a) meansfor deflecting the aim symbol vertically relative to a selected fixeddepression angle according to the error between measured altitude and aselected constant altitude; and (b) means for deflecting the aircraftsymbol vertically relative to a selected fixed depression angleaccording to a dynamic lead signal in order to prevent guidance controlovershoot and oscillation, whereby by maneuvering the aircraft so thatsaid aircraft symbol maintains the same depression angle as said aimsymbol the pilot will cause the aircraft to hold the selected altitude.11. The invention according to claim 10 further comprising:means forgenerating a signal equivalent to said selected fixed depression angle;means for generating a signal equivalent to the error in measuredaircraft altitude relative to the selected constant altitude; andsumming amplifier means for combining said selected fixed depressionangle signal with said error equivalent signal.
 12. The inventionaccording to claim 10 wherein signals corresponding to washed-out pitch,pitch rate, washed-out altitude rate and normal acceleration obtainedfrom aircraft motion longitudinal measurements are combined to providesaid dynamic lead signal.
 13. The invention according to claim 12wherein guidance is smoothly transitioned from said altitude hold modeto the descent mode by means of reducing the time constants of thewashed-out pitch and washed-out altitude rate during the period oftransition.
 14. The invention according to claim 13 furthercomprising:switching means to remove the altitude error signal from theaim symbol deflecting means when the pilot sees the aircraft symbolsuperimposed over the point on the ground toward which he wishes todescend, and to transfer the aim symbol deflecting means to saidselected fixed depression angle.
 15. The invention according to claim 13further comprising:switching means to remove the aim symbol from thedisplay when the pilot observes superposition of the aircraft symbolover his aimpoint, thereby removing the aim symbol as a guidancereference, the guidance reference being transferred to said aimpoint.16. The invention according to claim 13 further comprising:switchingmeans to convert the aim symbol to display a perspective view of anaimpoint controlled in response to a ground-based radio directional beamupon superposition of the aircraft symbol over the aimpoint.
 17. Theinvention according to claim 13 wherein said aircraft symbol providingaltitude hold guidance is converted to provide descent path guidancefurther comprising:switching means to reduce the time constants of thedynamic lead signal wash-out circuits when the pilot observessuperposition of the aircraft symbol over his aimpoint, thereby allowinga quicker change of said dynamic lead signal from the altitude holdsteady state value to an approach steady state value, and to increasethe time constants of the dynamic lead signal wash-out circuits to theiroptimum value for tracking the selected angle path to the aimpoint whenthe selected angle path has been acquired, thereby completing the smoothtransition to the descent mode.
 18. Apparatus for providing guidanceinformation to an aircraft pilot in the form of a conformal head-updisplay (HUD) which comprises:means for generating an aircraft symbol onthe display that simulates the attitude, location and motion of hisaircraft relative to the earth; means for generating an aim symbol thatis positioned on the display to provide a reference to which theaircraft symbol is directed for proper operation of the aircraft inselected flight modes; and means for generating symbols that simulate onthe display selected ground cues in position and orientation whichsuperimpose over actual selected ground cues as seen by the pilot whenthey are visible through the display; the simulating symbol generatingmeans comprising: (a) means for receiving radial directional beamsignals from a signal source located adjacent an airport runway, whichsource transmits a pattern of radio signals having a beam center andsignals distinguishing displacement from beam center within the pattern,the receiving means including means for detecting from the receivedradio beam signals the deviation of the aircraft relative to the beamcenter; (b) means responsive to the detecting means to locate in thedisplay corresponding to the pilot's perspective view of the runway asimulated runway including a runway center line and an aimpoint on thecenter line toward which to aim the landing approach; and furtherincluding (i) means for manually supplying a magnetic bearing signalindicative of the orientation of the centerline of the runway beingapproached; (ii) means for generating a magnetic heading signalindicative of the orientation of the aircraft; (iii) summing amplifiermeans for combining the magnetic bearing signal and the magnetic headingsignal to obtain a course error signal corresponding to the differencebetween the magnetic bearing signal and the magnetic heading signal;(iv) means for generating an eye height signal equivalent to the heightof the pilot's eyes relative to the terrain; (v) means for generating anoffset signal equivalent to the perpendicular displacement of thepilot's eyes from an extension of the runway centerline; (vi) divideramplifier means for generating a perspective ratio signal equivalent tothe ratio of said pilot's eye height to said pilot's eye displacement;(vii) means for deriving a vertical angle signal corresponding to thevertical displacement of the aimpoint relative to the horizontal; and(viii) variable gain amplifier means for varying the amplitude of saidvertical signal angle according to the amplitude of the perspectiveratio signal.
 19. The apparatus of claim 18 further including means forgenerating a glideslope error signal proportional to displacement of theaircraft from said beam center, means for developing signalscorresponding to the inastantaneous rate of change of said glideslopeerror signal, means for generating a fixed signal proportional to apreselected approach angle of descent, means for generating a leadcompensation signal, summing amplifier means for combining in a selectedratio said glideslope error and glideslope rate signals with said fixedsignal and said lead compensation signal for providing an amplifiedvertical deflection signal, and means for applying said amplifiedvertical deflection signal to the aircraft symbol generating means toincrease the separation between the aircraft symbol and the aim symbolfor a given glideslope error.
 20. Apparatus for providing guidanceinformation to an aircraft pilot in the form of a conformal head-updisplay (HUD) which comprises:means for generating an aircraft symbol onthe display that simulates the attitude, location and motion of hisaircraft relative to the earth; means for generating symbols thatsimulate on the display selected ground cues in position and orientationwhich superimpose over actual selected ground cues as seen by the pilotwhen they are visible through the display; means for generating an aimsymbol that is positioned on the display to provide a reference to whichthe aircraft symbol is directed for proper operation of the aircraft inselected flight modes; means for controlling the respective verticalpositions of the aircraft symbol and the aim symbol in response toselected signal inputs to provide vertical guidance information; meansfor receiving signals corresponding to a glideslope radio directionalbeam; means for generating a glideslope error signal corresponding toaircraft position relative to said radio directional beam; means fordeveloping signals corresponding to the instantaneous rate of change ofsaid glideslope error signal; aircraft instrumentation means forproviding signals corresponding to pitch angle, altitude rate and normalacceleration; means for generating a fixed signal proportional to apreselected approach angle of descent; means for generating a leadcompensation signal in accordance with said preselected angle of descentsignal and at least one of said signals from the instrumentation means;summing amplifier means for combining in a selected ratio saidglideslope error and glideslope rate signals with said fixed signal andsaid lead compensation signal to develop an aircraft symbol verticaldeflection signal; and means for applying said vertical deflectionsignal to the means for controlling the position of the aircraft symbolto develop an augmented displacement of the aircraft signal relative tothe aim signal for a given glideslope error.
 21. The apparatus of claim1 further including means for generating a fixed signal proportional toa preselected approach angle of descent, means for generating a leadcompensation signal, means for receiving a glideslope radio directionalbeam signal, means for generating a glideslope error signalcorresponding to aircraft position relative to the glideslope beam,summing amplifier means for combining in a selected ratio saidglideslope error and glideslope rate signals with said fixed signal andsaid lead compensation signal to provide an amplified signal indicativeof vertical displacement error, and means for applying said amplifiedsignal to the means for controlling the position of the aircraft signalin order to develop an augmented vertical displacement of the aircraftsymbol relative to the aim symbol for a given glideslope error.